1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to an air-cooled turbine rotor blade with a thick TBC and a low cooling flow.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a high temperature gas flow is passed through the turbine to produce mechanical work to drive the compressor and, in an industrial gas turbine engine, to also drive an electric generator and produce electrical energy. Passing a higher temperature gas flow into the turbine can increase the efficiency of the engine. However, the turbine inlet temperature is limited by the material properties of the first stage stator vanes and rotor blades as well as the amount of cooling that can be produced by passing cooling air through these airfoils (vanes and blades). If the turbine inlet temperature is too high, then the first stage vanes and blades can become too hot and even melt. Thus, one method of increasing the turbine inlet temperature is to form the turbine vanes and blades from even higher temperature resistant materials.
Another method of allowing for an increase in the turbine inlet temperature is to provide cooling for the airfoils. Airfoil designers try to minimize the amount of cooling air used in the airfoils since the cooling air is typically bled off from the compressor and thus is not used to produce work and the energy used to compress the air is thus wasted. Complex airfoil internal cooling circuits have been proposed that include combinations of convection cooling, impingement cooling and even film cooling of the airfoil outer surfaces.
FIG. 1 shows a typical first stage turbine blade external pressure profile. As seen in FIG. 1, the forward region of the pressure side surface experiences a higher hot gas static pressure while the entire suction side external surface of the airfoil is at a much lower hot gas static pressure than the pressure side. The vertical dashed line in FIG. 1 represents the highest pressure on the external surface of the airfoil just downstream from the leading edge region. One can see that the pressure on the suction side opposite from the highest pressure on the pressure side is much lower.
FIGS. 2 through 4 shows a prior art cooling circuit for a first stage turbine blade in an industrial gas turbine (IGT) engine. This cooling circuit is referred to as a 1+5+1 forward flowing serpentine cooling circuit and includes a leading edge cooling air supply channel 11 located in the leading edge region of the airfoil to supply cooling air to a leading edge impingement cavity 12 through a row of metering and impingement holes 13, and with a showerhead arrangement of film cooling holes 14 and gill holes 15 on both sides of the leading edge region to provide film cooling on the leading edge region.
The airfoil mid-chord region is cooled by a 5-pass forward flowing serpentine flow circuit that includes a first leg or channel 21 adjacent to a trailing edge region, followed by the second leg 22, third leg 23, fourth leg 24 and fifth leg 25 to form the serpentine flow path. As seen in FIG. 2, film cooling holes 35 are used on the pressure side and suction side walls to discharge cooling air from some of the legs 21-25 that form the serpentine flow circuit.
Also seen in FIGS. 2 and 4 is the trailing edge region cooling circuit that includes a trailing edge cooling air supply channel 31 that feeds into a row of metering and impingement holes 32 and impingement cavities 33 that form a series of metering and impingement holes followed by impingement cavities to provide cooling for the trailing edge region. A row of cooling air exit holes is arranged along the trailing edge to discharge the cooling air. A row of film cooling holes 35 is connected to the first impingement cavity 33 to discharge film cooling air onto the pressure sidewall.
For a forward flowing 5-pass serpentine cooling design of FIGS. 2-4 used in the airfoil mid-chord region, the cooling air flows toward the leading edge and discharges into the high hot gas side pressure section of the pressure side. In order to satisfy the back flow margin (the hot gas flow does not flow into the internal cooling passages of the airfoil), a high cooling air supply pressure is needed for the FIG. 2 design, and therefore will induce a high leakage flow. In the FIG. 2 airfoil cooling circuit, the blade tip section is cooled with two tip turns in conjunction with local film cooling. Cooling air bleed off from the 5-pass serpentine flow circuit will reduce the cooling performance for the serpentine flow circuit. Independent cooling flow circuits from the mid-chord cooling circuit is used to provide cooling for the airfoil leading and trailing edges.
As the TBC technology improves and more IGT engine turbine blades are applied with relatively thick or low conductivity TBC, the amount of cooling air required is reduced. As a result, there is not sufficient cooling airflow for the prior art 1+5+1 cooling circuit of FIGS. 2-4. Cooling air flow for the blade leading edge trailing edges has to be combined with the mid-chord cooling circuit to form a single 5-pass flow circuit in order to provide adequate cooling for the entire airfoil using the low flow cooling air used for low cooling flow airfoils. However, for a single forward flowing 5-pass serpentine cooling circuit with total blade cooling flow, the BFM (back flow margin) may become a serious design issue.
In order to avoid the BFM issue described above in the FIG. 2 cooling circuit, the forward flowing 5-pass serpentine circuit of FIG. 2 can be transformed into an aft flowing 5-pass serpentine circuit as seen in the FIGS. 5 and 6 design. The FIGS. 5 and 6 design transforms the airfoil cooling with a single 5-pass aft flowing serpentine cooling circuit that includes a forward section leading edge impingement cavity 46 and an aft flowing serpentine flow circuit with a first leg 41 located adjacent to the impingement cavity 46, a second leg 42, a third leg 43, a fourth leg 44 and a fifth leg 45 that forms the 5-pass serpentine aft flowing circuit. A row of metering and impingement holes 47 connects the first leg 41 to the impingement cavity 46, and a showerhead arrangement of film cooling holes 48 connects the impingement cavity 46 to discharge the layer of film cooling air onto the leading edge of the airfoil. The fifth leg 45 is connected to a row of trailing edge exit holes 49 to discharge the spent serpentine flow cooling air through the trailing edge of the airfoil.
For the forward section of the blade leading edge impingement cooling in the FIG. 5 designs, it is normally designed in conjunction with leading edge backside impingement cooling plus a showerhead arrangement of film cooling holes with pressure side and suction side film discharge cooling holes (not shown in FIG. 5 or 6). Cooling air is supplied from the first up-pass channel 41 of the 5-pass serpentine circuit. The impingement cooling air is normally fed through a row of metering holes 47, and impinged onto the backside of the airfoil leading edge surface to provide backside impingement cooling of the leading edge prior to discharging the spent impingement cooling air as film cooling air through the showerhead holes and the P/S and S/S gill holes. One possible drawback for the 5-pass aft flowing serpentine cooling circuit of FIGS. 5 and 6 is the heat pick up by the cooling flow. As the cooling air reaches the airfoil trailing edge, the heated cooling air looses its cooling potential since the cooling air is being heated as it travels through the 5 legs of the serpentine circuit. Thus, with the cooling circuit of FIGS. 5 and 6, a turbine upgrade may become a design limitation.